Low lump mass combustor wall with quench aperture(s)

ABSTRACT

An assembly is provided for a turbine engine. This turbine engine assembly includes a combustor wall including a first layer vertically connected with a second layer. A first portion of the first layer overlaps and is vertically spaced from the second layer by a cavity. A second portion of the first layer is substantially vertically inline with an adjacent portion of the second layer. The second portion of the first layer at least partially forms a quench aperture vertically through the combustor wall.

This application claims priority to U.S. Patent Appln. No. 62/074,911filed Nov. 4, 2014.

BACKGROUND OF THE INVENTION

1. Technical Field

This disclosure relates generally to a turbine engine and, moreparticularly, to a combustor for a turbine engine.

2. Background Information

A floating wall combustor for a turbine engine typically includes abulkhead, an inner combustor wall and an outer combustor wall. Thebulkhead extends radially between the inner and the outer combustorwalls. Each combustor wall may include a shell and a heat shield, whichheat shield forms a respective radial peripheral side of a combustionchamber. Cooling cavities extend radially between and separate the heatshield and the shell. These cooling cavities may fluidly coupleimpingement apertures in the shell with effusion apertures in the heatshield.

Each combustor wall may also include a plurality of quench aperturegrommets located between the shell and the heat shield. Each of thequench aperture grommets forms a quench aperture radially through therespective combustor wall. The quench aperture grommets as well asadjacent portions of the heat shield are typically subject to relativelyhigh temperatures during turbine engine operation due to the relativelylarge lump material mass associated therewith, which can inducerelatively high thermal stresses within the grommets and the heatshield.

There is a need in the art for an improved turbine engine combustor.

SUMMARY OF THE DISCLOSURE

According to an aspect of the invention, an assembly is provided for aturbine engine. This turbine engine assembly includes a combustor wallincluding a first layer vertically connected with a second layer. Afirst portion of the first layer overlaps and is vertically spaced fromthe second layer by a cavity. A second portion of the first layer issubstantially vertically inline with an adjacent portion of the secondlayer. The second portion of the first layer at least partially forms aquench aperture vertically through the combustor wall.

According to another aspect of the invention, a combustor wall isprovided for a turbine engine. This combustor wall includes a firstlayer vertically attached with a second layer. A first portion of thefirst layer overlaps and is vertically offset from the second layer. Asecond portion of the first layer extends vertically at least partiallyinto an aperture in the second layer. The second portion of the firstlayer at least partially forms a quench aperture vertically through thecombustor wall.

According to still another aspect of the invention, another combustorwall is provided for a turbine engine. This combustor wall includes afirst layer vertically connected with a second layer. A cooling cavityextends vertically between the first layer and the second layer. Thefirst layer substantially completely defines a quench aperturevertically through the combustor wall. The first layer has asubstantially uniform thickness.

The first layer may substantially completely define the quench aperturevertically through the combustor wall.

The second layer may have a substantially uniform thickness.

The first layer may include or be configured as a shell. The secondlayer may include or be configured as a heat shield. Alternatively, thefirst layer may include or be configured as a heat shield. The secondlayer may include or be configured as a shell.

The second portion may substantially completely define the quenchaperture vertically through the combustor wall.

A thickness of the first portion may be substantially equal to athickness of the second portion.

The second portion may extend at least partially vertically into anaperture in the second layer.

The second portion may form a vertical indentation in the first layer.

At least the second portion of the first layer and the second layer maybe configured to form a side periphery of a combustion chamber.

At least a portion of the first layer at the side periphery may becoated with a thermal barrier coating. In addition or alternatively, atleast a portion of the second layer at the side periphery may be coatedwith a thermal barrier coating.

The second portion may have a curved geometry.

The second portion may include one or more cooling holes.

The second portion may include one or more stiffening features.

The second portion may include one or more protrusions.

The combustor wall may extend circumferentially about and axially alonga centerline. The first portion may be located axially forward of thesecond portion. Alternatively, the first portion may be located axiallyaft of the second portion.

A third portion of the first layer may overlap and be vertically spacedfrom the second layer by the cavity or another cavity. The secondportion may be between the first and the third portions.

The foregoing features and the operation of the invention will becomemore apparent in light of the following description and the accompanyingdrawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a side cutaway illustration of a geared turbine engine.

FIG. 2 is a partial sectional schematic illustration of an assembly forthe turbine engine of FIG. 1.

FIG. 3 is a perspective schematic illustration of a combustor for theturbine engine assembly of FIG. 2.

FIG. 4 is a partial sectional schematic illustration of a combustor wallfor the turbine engine assembly of FIG. 2.

FIGS. 5-11 are partial sectional schematic illustrations of alternatecombustor walls for the turbine engine assembly of FIG. 2.

FIGS. 12 and 13 are partial perspective illustrations of a portion of acombustor wall with alternative stiffening and/or heat transferfeatures.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 is a side cutaway illustration of a geared turbine engine 20.This turbine engine 20 extends along an axial centerline 22 between anupstream airflow inlet 24 and a downstream airflow exhaust 26. Theturbine engine 20 includes a fan section 28, a compressor section 29, acombustor section 30 and a turbine section 31. The compressor section 29includes a low pressure compressor (LPC) section 29A and a high pressurecompressor (HPC) section 29B. The turbine section 31 includes a highpressure turbine (HPT) section 31A and a low pressure turbine (LPT)section 31B.

The engine sections 28-31 are arranged sequentially along the centerline22 within an engine housing 32. Each of the engine sections 28, 29A,29B, 31A and 31B includes a respective rotor 34-38. Each of these rotors34-38 includes a plurality of rotor blades arranged circumferentiallyaround and connected to one or more respective rotor disks. The rotorblades, for example, may be formed integral with or mechanicallyfastened, welded, brazed, adhered and/or otherwise attached to therespective rotor disk(s).

The fan rotor 34 is connected to a gear train 40, for example, through afan shaft 42. The gear train 40 and the LPC rotor 35 are connected toand driven by the LPT rotor 38 through a low speed shaft 43. The HPCrotor 36 is connected to and driven by the HPT rotor 37 through a highspeed shaft 44. The shafts 42-44 are respectively rotatably supported bya plurality of bearings 46; e.g., rolling element and/or thrustbearings. Each of these bearings 46 may be connected to the enginehousing 32 by at least one stationary structure such as, for example, anannular support strut.

During operation, air enters the turbine engine 20 through the airflowinlet 24, and is directed through the fan section 28 and into a core gaspath 48 and a bypass gas path 50. The air within the core gas path 48may be referred to as “core air”. The air within the bypass gas path 50may be referred to as “bypass air”. The core air is directed through theengine sections 29-31, and exits the turbine engine 20 through theairflow exhaust 26 to provide forward engine thrust. Within thecombustor section 30, fuel is injected into a (e.g., annular) combustionchamber 52 and mixed with the core air. This fuel-core air mixture isignited to power the turbine engine 20. The bypass air is directedthrough the bypass gas path 50 and out of the turbine engine 20 througha bypass nozzle 54 to provide additional forward engine thrust which mayaccount for the majority of the forward engine thrust. Alternatively, atleast some of the bypass air may be directed out of the turbine engine20 through a thrust reverser to provide reverse engine thrust.

FIG. 2 illustrates an assembly 56 of the turbine engine 20 of FIG. 1.The turbine engine assembly 56 includes a combustor 58 arranged within a(e.g., annular) combustor plenum 60 formed by a diffuser module. Theplenum 60 receives compressed core air from the HPC section 29B throughan inlet passage of the diffuser module. The plenum 60 provides thereceived core air to the combustor 58 as described below in furtherdetail.

The turbine engine assembly 56 also includes one or more fuel injectorassemblies 62 arranged circumferentially around the centerline 22. Eachof these fuel injector assemblies 62 includes a fuel injector 64 thatmay be mated with a swirler 66. The fuel injectors 64 inject the fuelinto the combustion chamber 52. The swirlers 66 direct some of the coreair from the plenum 60 into the combustion chamber 52 in a manner thatfacilitates mixing the core air with the injected fuel. One or moreigniters (not shown) ignite the fuel-core air mixture. Quench apertures68 (see also FIG. 3) in inner and/or outer walls 70 and 72 of thecombustor 58 may direct additional core air into the combustion chamber52 for combustion. Additional core air may also be directed (e.g.,effused) into the combustion chamber 52 through one or more coolingholes (e.g., see 74-76 in FIGS. 4 and 9) in the inner and the outerwalls 70 and 72.

The combustor 58 may be configured as an annular floating wallcombustor. The combustor 58 of FIG. 2, for example, includes an annularcombustor bulkhead 78, the tubular combustor inner wall 70, and thetubular combustor outer wall 72. The bulkhead 78 extends radiallybetween and is connected to the inner wall 70 and the outer wall 72.Each wall 70, 72 extends axially downstream and aft along the centerline22 from the bulkhead 78 towards the HPT section 31A, thereby definingthe combustion chamber 52.

Each combustor component 70, 72 and 78 may be a multi-walled structurethat includes, for example, an interior layer (e.g., a heat shield 80)connected with an exterior layer (e.g., a shell 82). The inner and theouter walls 70 and 72, for example, each respectively include a heatshield 80 attached (e.g., mechanically fastened and/or bonded) to ashell 82 with at least one cooling cavity 84 (e.g., impingement cavity)extending vertically (e.g., generally radially relative to thecenterline 22) between the shell 82 and the heat shield 80. Referring toFIG. 4, this cooling cavity 84 may be fluidly coupled with the plenum 60through the one or more impingement holes 74 in the shell 82. Thecooling cavity 84 may be fluidly coupled with the combustion chamber 52through the one or more effusion holes 75 (see also 76 in FIG. 9) in theheat shield 80 and/or shell 82. The shell 82 may be configured as aunitary full hoop body. The heat shield 80 may also be configured as aunitary full hoop body. The present disclosure, however, is not limitedto the foregoing multi-walled structure configuration. For example, insome embodiments the shell 82 may be configured from one or more arraysof shell panels. In addition or alternatively, in some embodiments, theheat shield 80 may be configured from one or more arrays of heat shieldpanels.

Referring to FIGS. 2-4, the shell 82 of the inner and/or the outer walls70 and 72 may include one or more vertical indentations 86; e.g.,depressions and/or cupped regions. Each of these indentations 86 may beconfigured as a portion 88 of the shell 82 which is vertically depressed(e.g., recessed radially inward or outward) relative to one or moreadjacent and/or surrounding portions 90 and 92 of the shell 82. Forexample, referring to FIG. 4, an exterior surface 94 (e.g., a plenumsurface) of the shell 82 at (e.g., on, adjacent or proximate) theportion 88 is vertically offset from the exterior surface 94 at theadjacent portions 90 and 92. A (e.g., annular) transition 96 between theportions may be gradual (e.g., sloped) as illustrated in FIG. 4, oralternatively sharp as illustrated in FIG. 5.

Each portion 88 is configured to substantially completely (or at leastpartially) define a respective one of the quench apertures 68 throughthe combustor wall 70, 72. The portion 88 of FIG. 4, for example,extends vertically at least partially into (or through) a respectiveaperture 98 in the heat shield 80. In this manner, the combustor wall70, 72 at the respective quench aperture 68 may have a single layerstructure; e.g., a shell only structure. It is worth noting, such asingle layer structure may significantly reduce the material lump massproximate the quench aperture 68 as compared to prior art combustorwalls which include grommets and/or other relatively thick annularbodies to define quench apertures. As a result, the combustor wall 70,72 of the present disclosure may be subject to lower thermal stressesand fatigues than prior art combustor walls.

Referring again to FIG. 4, the portion 88 may be substantiallyvertically inline with an adjacent portion 100 of the heat shield 80which defines the aperture 98. Alternatively, the portion 88 may bevertically recessed within the aperture 98 (see FIG. 6), or protrudevertically out from the aperture 98 into the combustion chamber 52 (seeFIG. 7). With the foregoing configurations, the portion(s) 88 and theheat shield 80 together form a side periphery of the combustion chamber52. Thus, to provide a protection between combustion gases within thecombustion chamber 52 and the combustor wall 70, 72, each portion 88and/or the heat shield 80 (or at least a portion thereof) as well as, insome embodiments, the transition(s) 96 may be coated with a coating suchas, but not limited to, a thermal barrier coating. Various types ofcoatings and thermal barrier coatings are known in the art and thepresent disclosure is not limited to any particular types thereof. It isalso worth noting, the coating may be relatively thin since the portionsof the combustor wall 70, 72 proximate the quench aperture may bemaintained relatively cool.

An outer peripheral geometry (e.g., size and/or shape) of the portion 88may be similar, but slightly smaller than an inner peripheral geometryof the aperture 98. In this manner, a channel 102 (e.g., an annularcooling channel) may be formed between the portions 88 and 100. Thischannel 102 fluidly couples the cooling cavity 84 with the combustionchamber 52. Alternatively, referring to FIG. 8, the portions 88 and 100may fully (or partially) engage one another to provide a substantiallysealed interface between the shell 82 and the heat shield 80.

The outer peripheral geometry of the portion 88 may be curved (e.g.,circular, oval, elliptical, etc.) as illustrated in FIG. 3, polygonal(e.g., square, rectangular, triangular, etc.) or any other shape. Thepresent disclosure, of course, is not limited to any particular outerperipheral geometry shapes.

One or more of the portions 90 and 92 may each have a vertical thicknessthat is substantially equal to a vertical thickness of the portion 88.In such an embodiment, the shell 82 may have a substantially uniformthickness. Alternatively, one or more of the portions 90 and 92 may eachhave a vertical thickness that is less than or greater than a verticalthickness of the portion 88.

Each of the portions 90 and 92 may overlap and be verticallyspaced/offset from the respective heat shield 80. In this manner, theportions 90 and 92 and respective vertically opposing portion(s) 100 ofthe heat shield 80 may form the cooling cavity 84 in the combustor wall70, 72.

In some embodiments, as illustrated in FIG. 9, each portion 88 mayinclude the one or more cooling holes 76; e.g., effusion apertures.These cooling holes 76 may be arranged circumferentially about therespective quench aperture 68. Each cooling hole 76 of FIG. 9 extendsvertically through an annular rim 104 of the portion 88, which rim 104forms an outer periphery of the respective quench aperture 68. It isworth noting, the cooling holes 76 receive cooling air from the plenum60 opposed to the cooling holes 75 which receive cooling air from thecavity 84. As a result, each portion 88 of the shell 82 can receive ahigher degree of film cooling than the adjacent portion(s) 100 of theheat shield 80.

In some embodiments, as illustrated in FIG. 10, the heat shield 80 ofthe combustor wall 70, 72 may be configured forward of the portions 88and the quench apertures 68. In such embodiments, the portions 88 and 92may be configured together as a single portion that extends to an aft,downstream distal end of the combustor wall 70, 72. In other embodimentshowever, as illustrated in FIG. 11, the heat shield 80 of the combustorwall 70, 72 may be configured aft of the portions 88 and the quenchapertures 68. In such embodiments, the portions 88 and 90 may beconfigured together as a single portion that extends to a forward,upstream distal end of the combustor wall 70, 72.

In some embodiments, as illustrated in FIGS. 12 and 13, one or more ofthe portions 88 (or any other portion of the shell 82 and/or heat shield80) may each be configured with one or more stiffening and/or heattransfer features 106. Examples of a stiffening and/or heat transferfeature 106 include, but are not limited to, a rib/channel (see FIG.12), a point protrusion (see FIG. 13), a dimple and/or any other type ofprojection and/or recession. Such features 106 may be configured toincrease the structural rigidity of the portion(s) 88. Such features 106may also or alternatively be configured to increase convective heattransfer between the shell 82 and the gas within the plenum 60.

In some embodiments, the combustor wall 70, 72 may include one or morerails. These rails may be configured with (e.g., included with orattached to) the shell 82 and/or the heat shield 80. Each rail mayextend vertically between the shell 82 and the heat shield 80 andthereby sub-divide the cooling cavity 84 into a plurality of coolingcavities. For example, one or more of the rails may be aligned with oneor more of the portions 88 such that each portion 88 is between andadjacent two or more different cavities.

In some embodiments, one or more of the combustor walls 70 and 72 mayalso be configured with one or more traditional quench apertures; e.g.,quench apertures formed by quench aperture grommets. For example, thecombustor wall 70, 72 may be configured with a mixture of alternatingtraditional quench apertures and quench apertures 68 formed as describedabove. In other embodiments, one of the combustor walls 70, 72 may beconfigured as described above while the other one of the combustor walls72, 70 may have a more traditional configuration; e.g., quench aperturesformed by quench aperture grommets.

The turbine engine assembly 56 may be included in various turbineengines other than the one described above. The turbine engine assembly56, for example, may be included in a geared turbine engine where a geartrain connects one or more shafts to one or more rotors in a fansection, a compressor section and/or any other engine section.Alternatively, the turbine engine assembly 56 may be included in aturbine engine configured without a gear train. The turbine engineassembly 56 may be included in a geared or non-geared turbine engineconfigured with a single spool, with two spools (e.g., see FIG. 1), orwith more than two spools. The turbine engine may be configured as aturbofan engine, a turbojet engine, a propfan engine, or any other typeof turbine engine. The present invention therefore is not limited to anyparticular types or configurations of turbine engines.

While various embodiments of the present invention have been disclosed,it will be apparent to those of ordinary skill in the art that many moreembodiments and implementations are possible within the scope of theinvention. For example, the present invention as described hereinincludes several aspects and embodiments that include particularfeatures. Although these features may be described individually, it iswithin the scope of the present invention that some or all of thesefeatures may be combined with any one of the aspects and remain withinthe scope of the invention. Accordingly, the present invention is not tobe restricted except in light of the attached claims and theirequivalents.

What is claimed is:
 1. An assembly for a turbine engine, comprising: acombustor wall including a first layer vertically connected with asecond layer; a first portion of the first layer overlapping andvertically spaced from the second layer by a cavity; and a secondportion of the first layer vertically inline with an adjacent portion ofthe second layer, and at least partially forming a quench aperturevertically through the combustor wall; wherein a vertical thickness ofthe first portion is equal to a vertical thickness of the secondportion; and wherein the combustor wall extends circumferentially aboutand axially along a centerline, and the first portion is located axiallyforward of the second portion.
 2. The assembly of claim 1, wherein thesecond portion completely defines the quench aperture vertically throughthe combustor wall.
 3. The assembly of claim 1, wherein the secondportion extends partially vertically into an aperture in the secondlayer.
 4. The assembly of claim 1, wherein the second portion forms avertical indentation in the first layer.
 5. The assembly of claim 1,wherein at least the second portion of the first layer and the secondlayer are configured to form a side periphery of a combustion chamber.6. The assembly of claim 5, wherein at least a portion of the firstlayer at the side periphery is coated with a thermal barrier coatingand/or at least a portion of the second layer at the side periphery iscoated with a thermal barrier coating.
 7. The assembly of claim 1,wherein the second portion has a curved geometry.
 8. The assembly ofclaim 1, wherein the second portion includes one or more cooling holes.9. The assembly of claim 1, wherein the second portion includes one ormore stiffening features.
 10. The assembly of claim 1, wherein thesecond portion includes one or more protrusions.
 11. The assembly ofclaim 1, wherein the first layer comprises a shell and the second layercomprises a heat shield.
 12. An assembly for a turbine engine,comprising: a combustor wall including a first layer verticallyconnected with a second layer; a first portion of the first layeroverlapping and vertically spaced from the second layer by a cavity; anda second portion of the first layer vertically inline with an adjacentportion of the second layer, and at least partially forming a quenchaperture vertically through the combustor wall; wherein a verticalthickness of the first portion is equal to a vertical thickness of thesecond portion; and wherein the combustor wall extends circumferentiallyabout and axially along a centerline, and the first portion is locatedaxially aft of the second portion.
 13. An assembly for a turbine engine,comprising: a combustor wall including a first layer verticallyconnected with a second layer; a first portion of the first layeroverlapping and vertically spaced from the second layer by a cavity; anda second portion of the first layer vertically inline with an adjacentportion of the second layer, and at least partially forming a quenchaperture vertically through the combustor wall; wherein a verticalthickness of the first portion is equal to a vertical thickness of thesecond portion; and wherein a third portion of the first layer overlapsand is vertically spaced from the second layer by the cavity or anothercavity, and the second portion is between the first portion and thethird portion.